The present invention relates generally to gas turbine engines, and, more specifically, to cooled turbine blades and vanes therein.
In a typical gas turbine engine, one or more stages of stationary turbine vanes and rotating turbine blades are disposed downstream of an annular combustor which discharges hot combustion gases from which energy is extracted by the rotor blades and suitably used for producing work. Since the high pressure turbine (HPT) rotor blades are disposed closest to the combustor, they are subject to the hottest combustion gas temperature and therefore typically include cooling circuits therein for maintaining the maximum temperature thereof within acceptable limits for obtaining a suitable useful life of the blades. The cooling circuits are passages or channels formed inside the airfoil portion of the blade by conventional casting techniques which carry air bled from the compressor of the engine for cooling the blade. As the air passes inside the blade it removes heat from the blade, with the cooling circuits typically being configured for maximizing the amount of heat removal for minimizing overall efficiency losses in the engine. Since any air bled from the compressor is not therefore being used in the combustion process for generating energy, the bleed air provides a performance penalty.
Accordingly, typical blade cooling circuits include conventional serpentine channels both for stator vanes and rotor blades which repeatedly channel cooling air outwardly and inwardly along the radial or longitudinal axis of the blades and vanes for removing a maximum amount of heat therefrom.
Turbine blades and vanes have airfoil portions which are generally crescent in configuration with opposite generally convex suction and generally concave pressure sides joined together along leading and trailing edges of the airfoil. Accordingly, the pressure and velocity profiles of the combustion gases which flow over the airfoil pressure and suction sides varies from the leading edge to the trailing edge of the airfoil. This, in turn, affects the temperature distribution over the entire surface of the airfoil from the leading edge to the trailing edge, with the temperature distribution also varying radially from the root to the tip of the airfoil as is conventionally known.
Accordingly, the cooling circuits inside the airfoil are typically designed for each application and the associated temperature and heat loads experienced by the airfoil over its outer surface. In addition to the different temperature environment experienced by the pressure and suction sides of the airfoil, the typical airfoil also has different temperature environments, and therefore cooling needs, at its leading edge region, mid-chord region, and trailing edge region. The cooling circuits within the airfoil are therefore typically tailored for each of these three regions as well as for the pressure and suction sides of the airfoil.
Various types of conventional cooling arrangements are well known in the art and include convection cooling, impingement cooling, and film cooling which are selectively used in blade and vane cooling designs for obtaining enhanced cooling thereof. The cooling air channeled inside the airfoil removes heat by convection as well as by impingement cooling therein in some designs. The spent cooling air is then discharged from the airfoil typically through the tip thereof as well as through the pressure and/or suction side as required. In the latter case, discharge holes are conventionally formed through the airfoil sides for discharging the cooling air in a film along the surface of the airfoil to provide an insulating film cooling barrier with the combustion gases flowable thereover. Film cooling holes are typically radially spaced apart from each other in columns extending between the airfoil root and tip and at selected axial locations between the airfoil leading and trailing edges. Film cooling has a limited axial duration, and therefore, axially spaced apart columns of film cooling holes are typically utilized as required to reestablish film cooling in the axial downstream direction along the airfoil.
Fundamental to effective film cooling is the conventionally known blowing ratio which is merely the product of the density and velocity of the discharge flow from the film cooling holes relative to the product of the density and velocity of the combustion gases at the outlets of the film cooling holes. Excessive blowing ratios cause the discharged cooling air to separate or blow-off from the airfoil outer surface which degrades film cooling effectiveness. Accordingly, the airfoil must be designed to ensure effective blowing ratios while minimizing blow-off tendency and preventing backflow of combustion gases through the film cooling holes into the blade. Since the pressure and velocity of the combustion gases flowing over the pressure and suction sides of the airfoil varies, multiple cooling circuits are typically provided through the airfoil to ensure that blowing ratios for each circuit are within acceptable minimum and maximum values to prevent backflow and blow-off, respectively.
Since significant differences in static pressures and velocities of the combustion gas flow between the pressure and suction sides of an airfoil exist, the blowing ratio of the film cooling air on the pressure side is usually much higher than that on the suction side when the film cooling holes are fed by a common cooling circuit within the airfoil which must be suitably accommodated for preventing film blow-off in the airfoil outer surface.